Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through formulation of iron, nickel and cobalt-based superalloys, though components formed from such alloys often cannot withstand long service exposures if located in certain sections of a gas turbine engine, such as the turbine, combustor or augmentor.
A common solution is to provide internal cooling of turbine, combustor and augmentor components, at times in combination with a thermal barrier coating. Airfoils of gas turbine engine blades and vanes often require a complex cooling scheme in which cooling air flows through cooling channels within the airfoil and is then discharged through carefully configured cooling holes at the airfoil surface. Convection cooling occurs within the airfoil from heat transfer to the cooling air as it flows through the cooling channels. In addition, fine internal orifices can be provided to direct cooling air flow directly against inner surfaces of the airfoil to achieve what is referred to as impingement cooling, while film cooling is often accomplished at the airfoil surface by configuring the cooling holes to discharge the cooling air flow across the airfoil surface so that the surface is protected from direct contact with the surrounding hot gases within the engine.
In the past, cooling channels have typically been integrally formed with the airfoil casting using relatively complicated cores and casting techniques. More recently, U.S. Pat. Nos. 5,626,462 and 5,640,767, both to Jackson et al. and commonly assigned with the present invention, teach a method of forming a double-walled airfoil by depositing an airfoil skin over a separately-formed inner support wall (e.g., a spar) having surface grooves filled with a sacrificial material. After the airfoil skin is formed, preferably by such methods as electron-beam physical vapor deposition (EBPVD) or plasma spraying, the sacrificial material is removed to yield a double-walled airfoil with cooling channels that circulate cooling air against the interior surface of the airfoil skin.
The sacrificial material must be carefully selected to withstand the skin deposition temperature, typically about 1000 to 1200.degree. C., and have adequate thermal conductivity to prevent overheating of the deposition (spar) surface. In addition, the sacrificial material must be deposited to completely fill the surface grooves of the spar, and have a coefficient of thermal expansion (CTE) that is at least as high as that of the spar to prevent shrinkage of the sacrificial material away from the groove wall during deposition of the airfoil skin. If incomplete fill or shrinkage occurs, a gap will be present between the sacrificial material and the groove wall during skin deposition, which, if sufficiently large, leads to an unacceptable surface defect in the airfoil skin. Airfoil skins deposited by EBPVD are particularly sensitive to surface discontinuities due to the atom-by-atom manner in which the coating is built up. For example, gaps of less than up to about 2 mils (about 50 Fm) can be tolerated by plasma-sprayed airfoil skins, while gaps should not be larger than about 0.5 mil (about 13 Fm) for airfoil skins deposited by EBPVD.
For the above reason, sacrificial materials in the form of a solid shaped to fit a surface groove of a spar are unacceptable, because of the inherent likelihood of a gap between the solid and the groove wall. Consequently, and as shown in FIGS. 1 through 3, Jackson et al. teach that a sacrificial material 12 is deposited in the surface groove 14 of a spar 10 in excess amounts, with the excess being removed by machining or another suitable technique so that the surface of the sacrificial material 12 is flush with the surrounding surface of the spar 10. Depending on its composition, the sacrificial material 12 can be removed after deposition of the airfoil skin 16 by melting/extraction, chemical etching, pyrolysis or another suitable method.
As is evident from FIG. 3, though a sacrificial material 12 having the desired physical properties outlined above can be deposited to completely fill the groove 14, a problem arises because the groove 14 inevitably has rounded corners and edges, particularly if formed by casting or machining. As a result, even if the sacrificial material 12 completely fills the groove 14 and does not shrink to form a gap during deposition of the airfoil skin 16, removal of the material 12 leaves a sharp (&lt;90.degree.) notch 18 between the skin 16 and the wall of the groove 14. Cracks will tend to initiate from this notch 18, and thereafter propagate between the airfoil skin 16 and spar 10.
In view of the above, it can be seen that it would be desirable if a method were available that could ensure adequate filling of the groove prior to deposition of the airfoil skin, while also eliminating notches that could serve as the source of cracks between the skin and spar.